# ADVISORY



# CIFtcULAR DEPARTMENT

OF TRANSPORTATION FEDERAL AVIATION ADMINISTRATION SUBJECT:

FATIGUE EVALUATIONOFROTORCRAFTSTKIClsJRE

10 

### PURPOSE.

This advisory circular sets forth acceptable meansof compliancewith the provisions of FederalAviation Regulations, Sections 27.571and29.571dealing with the fatigue evaluation of rotorcraft structure. 2 l 

### CANCELLATION.

AppendixA, MainRotor Service Life Determination, January 15, 1963, 

### to CAM 6is cancelled. 30 

### HOW TOGETTHISPUBLICATION.
a. Order copies of this publication from:

Departmentof Transportation FederalAviation Administration Distribution Unit, TAD 443.1 

### Washington, D.C. 20591 b

. Identify the publication in your order as: Fatigue Evaluation of Rotorcraft Structure, ACNo. 2w5.
C. 

### This advisory circular will be distributed free of charge.
‘1 I L Li ( 7g. ‘; 76’ t,;r ’ tJ. A. FERRARESE ‘Acting Director, Flight Standards Service

Initiated by: A&120AC20-95

SECTIONlo ZNTRODUCTION 1

-. . 10 Background

. 1 SECTIONI?

0

tXCTION 3.

XXTION. 5. EXtI,-SAFE ST~EMGTH EVALUATION 10 5/=/76

TABLEOFCONTENTS

2 l General 2 3l Instrumentation 3 4l Parts to be strain-gaged 3 5l Flight regimes and conditions to be investigated 3 l7XQUENCYOFLOADING 6l <Typeof operation 70 Loading spectrum FATIGUESTR.ENGTf!

EVALUATION 8 l General 90 Analytical method 100 Testing methods 11l Combination replacement time and fail-safe evaluation 12. Extension of replacement time

13 14: General Identification of critical portions of primary structure 150 Jlxtcnt of fail-safe damage 16l Determination of probable failure locations 170 Fail-safe clemonstration 180 Inspection Percent of occurrence 10 Power On- Rotor R.P.M. Airspeed Envelope 20 Power Off - Rotor R.P.M. Airspeed Envelope 3 4: Oscillatory Stress vs. Steady Stress Stress vs. Numberof Cycles 4 4 5 5 5 6 7 9 9

10 11 11. 1' L 12 13 14 16 16 16 16SECTION 1. INTRODUCTION

1 l 

### BACFRWND. The fatigue evaluation procedures outlined in this advisory circular are for guidance purposes only and are not mandatory nor regulatory in nature. Although auniform approach 

### to-.fatigue evaluation is desirable, it is recognized that in such a complex problem, new design features and methods of fabrication, new approaches to fatigue evaluation, and new configurations may require variations and deviations from the procedures described herein. Engineering judgment should therefore be exercised for each particular application.

a. The flight structure of the rotorcraft is subject to vibratory stresses in practically every regime of flight. In addition, since it is a highly maneuverable aircraft that is capable of forward, rearward, sideward, vertical, and rotational flight, operating limitations due to fatigue are possible in practically all flight situations. For these reasons, it is required that special attention be focused on the fatigue evaluation of the flight structure of the rotorcraft. b. Fatigue evaluation of the flight structure is intended to verify structural reliability. Assurance of strue tural rel iabili ty starts with design, including choice of materials for resistance to crack initiation and/or propagation, detail design to minimize stress concentration, and specification of surface finishes, fits, etc. Design analysis should include estimation of expected flight loads, and estimation of resistance to fatigue. Fatigue strength should be based on past full scale fatigue tests and/or materials fatigue data with appropriate reductions for the variability in fatigue strength, size, shape, surface finish, and environments of the structure. In addition, design for fatigue should consider mode-of-failure analysis, areas susceptible to fatigue cracking, and methods to assure detectability of fatigue cracks. The residual strength of a cracked structure is an important consideration of fail-safe design.

c. Assurance of structural reliability also includes manufacture and fabrication in accordance with design requirements and specifications, quality control to monitor compliance, and effective service inspection procedures.

Par 1 Page 1AC20-95 5/18/w

d l Fatigue evaluation of the structure,

. measurement of flight loads and stresses, and evaluation of fatigue strength and/or fatigue crack propagation are the subjects of this advisory circular. There is somequestion whether a cotipletely reliable method for the prediction of time to fatigue crack initiation and fracture exists. Nevertheless, one engineering approach to the subject is to use the "Linear Cumulative DamageHypothesis." This hypothesis states that every cycle of stress above an "endurance limit" produces fatigue damageproportional to the ratio of cycles accumulated at the stress to fatigue "life" at that stress,
e. Laboratory tests of this hypothesis indicate that it is reasonably valid when the loading spectrum consists of stresses which are, in effect, random6 Despite the lack of an adequate theory connecting this hypothesis with more basic properties of materials, it has been successfully used in a number of applications. f0 In addition, fatigue evaluation generally requires a method of accounting for the effect of steady loads and stresses on fatigue. Wherethe manufacturer does not provide other substantiating data, a Goodmandiagram maybe used to account for these effects. F:* In any rational fatigue evaluation, the following factors should be considered: (1) Identification of the structure to be considered in the fatigue evaluation. (2) The stresses associated with steady and maneuverinp operating I conditions expected in service. (3) The frequency of occurrences associated with various flight conditions and the corresponding spectrum of loadings and stresses. (4) The fatigue strength, fatigue crack propagation characteristics of the structure, and the residual strength of the cracked structure. SECTION2. FLIGHT STRAINMEASUREMENT PROGWU4 -

l l20 GENERAL. Subsequent to design analysis, in which aircraft loads and associated stresses are derived, the stress level and/or loads are to be verified by a carefully controlled flight strain measurement program.

Page 2 Par 1

### 5/18/N AC X1-95



### 30 INSTRUMENTATION.
a. The instrumentat-Len system used in the flight strain measurement program should accurately measure and record the critical strains under test conditions associated with normal operaticn and specific maneuvers. The location and distribution of the strain gages should be based on a rational evaluation of the critical stress areas. This maybe accomplished by appropriate analytical means, supplemented, whendeemednecessary, by strain sensitive coatings or photoelastic methods. The distribution and number of strain 

### egeS IlKlst def part essent i.a1 to theine sa the loadpspectrum ade

### qua tely for each fe operat ion of the rotorcraft.

p 0 The corresponding flight parameters (airspeed, rotor r.p.m., center of gravity accelerations, etc.) should also be recorded simultaneously by appropriate methods. This is necessary in order to correlate the loads and stresses with the maneuver or operating condition at which they occurred.

c. The instrumentation system should be adequately calibrated and checked periodically throughout the flight strain measurement program in order to ensure consistent and accurate results. 

### 4a PARTSTo

### BESTRAl+GAGED.Fatigue critical portions of the rotor systems, control systems, fuselage, and supporting structure for rotors, transmission, and engine are to be strain-gaged. For rotorcraft of unusual or unique design, special consideration might be necessary to insure that all of the essential parts are evaluated. s 5. 

### FLIGHT REGIMES

ANDCONDITIONS

TOBEINVXSTIGATED.
il. Typical flight and ground conditions to be investigated in the flipht.strain measurement program are given in Table 

### 1. Flight regimes that should be investigated for Dower-on andpower-off operation are shown in Figures 1 and 2. -For clarity, the parameters which define these regimes are included in these figures. As noted on Figure 1, complete coverage at 111 percent VNEshould be demonstrated for power-on operation. However, for power-off operation, Figure 2, complete coverage at 111percent VI\IEfor maximumand minimum design r.p.m. need not be obtained if points are obtained at Vm at both maximumand minimum design r.p.m. and at 111percent V~;Eat both maximumand minimum placarded r.p.m. as indicated in the figure.

Par 3 Page 3AC 20-95

b0 The determination of flight conditions to be investigated in the 'flight strain measurement program should be based on the antici- '

. pa&d use of the helicopter, and, if available, on past service records for similar designs. In any event, the flight conditions considered appropriate for the design and application should be representative of the actual operation in accordance with the rotorcraft flight manual. In the case of multiengine helicopters, the. flight conditions concerning partial engine-out operation should be considered in addition to complete power-off operation. The flight conditions to be investigated should be submitted in connection with the flight evaluation program. Suggested flight ., conditions for single-engine helicopters used in normal operation are shownin Table 1.

c. The severity of the maneuvers investigated during the flight strain survey should be such that it is unlikely that service use will be more severe. d l All flight conditions considered appropriate for the-particular design are to be investigated over the complete rotor speed, airspeed, center of gravity, altitude, andweight ranges in order to determine the mst critical stress levels associated with each flight condition. In order to account for data scatter and to determine the stress levels present, a sufficient amount of data , points should be obtained at each flight condition. In some instances, the critical weight, center of gravity, and altitude ranges for the various maneuvers can be based on past experience with similar designs. This procedure is acceptable where adequate flight tests are performed to substantiate such selections. The combinations of flight parameters that produce the most critical stress levels should be included in the fatigue evaluation.

SECTION3. FREQUENCY OFLOADIIJC 6 l TYPEOFOPERATION. The probable types of operation (transport, utility, etc.) for the rotorcraft should be established. The type of operation can have a major influence on the loading environment. 13ormally, the rotorcraft should be substantiated for the most critical general type of operation with consideration of special occasional types of operation.

Page4 , Par 5AC 20-95

70 LQkmGa The spectrum 

### allocating percentages of time or frequencies of occurrence to flight 

### conditions or maneuvers is to be baked on the expected usage of the rotorcraft. This spectrum is tosbe such that it is unlikly that actual usage will subject the structure to damagebeyond that associated with the spectrum. Considerations to be included in developing this spectrum should include 

### prior knowledgebasedon 

### flight history recorder data, 

### design limitations established in compliance with FAR’s 27,309 or 29.309 and recommended

### operating conditions and limitations specified 

### in the.rotorcraft flight manual. The distribution of times at various forward flight speeds should reflect not only the relation of these speeds to VNEbut also the recommended operating conditions in the rotorcraft flight manual which govern VCor 

### cruise speed.. Wherepossible, it is desirable to conduct the flight strain-gage 

### programbysimulating the usageas determined 

### above,with continuous recording of stresses and loads, thus 

### obtaining directly the stress/load spectra for structural elements. Table 1 contains typical percent of occurrences for the various .flight condixions for single-engine piston helicopters used in utility oper tions. This table should 

### beusedonly as a guide and should be modified as necessary for each particular rotorcraft.

m FATIGUESTRENGTH EVALUATION 80 

### GmERAL. Information to guide fatigue evaluation based on safe-life considerations leading to recommended replacement times is provided in this section. Although there is a large quantity of information 

### available onthe fatigue strength characteristics of material specimens, built-up specimens andparts, the prediction of strength of parts of newdesigns based on this information is less reliable than testing the actual part. Consequently, additional conservatism should be used with this method. However, in manycases the differences between past test specimens and the actual part (which 

### involve suchfactors as stress concentration, size, and fretting) cannot be accounted for with a reasonable degree of accuracy. Therefore, it is usually necessary. that the structural components be subjected to repeated load tests using information determined in the flight strain measurement program. Special operational or functional characteristics which could affect the fatigue strength should also be considered in the service life evaluation. Such factors as high blade operating temperatures due to tip jets or

Par 7 Page 5AC 

### 20-95 

## 5/18/N

turbine exhaust impingement 

### onthe tail rotor shouldbeconsidered aswell asother special operating conditions. 

### In addition, effects of special 

### purposeuse suchashoist and sling operation, spraying, surveying, etc. v 

### s.houldbeconsidered if appropriate to the particular type. Thefatigue strength shouldbeevaluated by either of the methodsoutlined below, but fYull scale testing is considered moreaccurate. 90 

### ANALY'TICPFL

METHOD,.

It is recognized that if allowab,ti bzess levels are established by acceptable means*andthe stresses measuredin flight are lower than these established levels, no fatigue testing is necessary. a. Simplified method. Thefollowing techniwes, basedonthe use of the Goodman

diagram,are

### considered acceptable for establishing this allowable stress level: Estimate the meanendurancelimit of the part from test results of specimenswith similar stress concentrations. The test specimenmaterial shouldberepresentative of the actual part andsufficient test data shouldbeavailable to substantiate the meanendurancelimit. Theestimate should accountfor surface conditions, fabrication methods,fretting, size andshapeeffects, aswell asdifferences in stress concentrations betweenthe test specimenandthe actual part. 

### Referring to figure 3, the meanendurancelimit mayberepresented by a straight 

### line drawnthrough the yield stress (point A on

### the horizontal. tis) andthe maximm 

### oscillatory stress whichthe aver especimencanwithstan at agive steady stress (Point B without failure for 10

7 ff 9 to 5x 10 

### cycles dependingonthe material. Afactor of safety of 3 shouldthen be applied to the mean endurancelimit sothat the slope of line ACwould be l/3 

### of line AB. Asmaller factor is acceptable whensubstantiated by asufficient numberof tests onsimilar parts in similar applications. If the flight strain measurementsindicate that al of apeating stresses fall belowthe operating boundaryline (AC), 

### nofatigue testing is necessary. When the measuredstresses .are abovethe operating boundaryline, however, fatigue testing of the actual parts shouldbe conducted. Caution shouldbeexercised in the application of the analytic& methodabove, particularly whenthe following items are involved:

Page 6 Par 85/18/7h AC m-95



## ( 1 a b) 

## ( 1 C 

### (a) ( > e . 

### Largeparts in proportion to the laboratory specimens. Irregularly shapedparts containing numerousor supeimposedfillets, holes, threads, or lugs. -. Parts of unique designfor which nopast service experience is available. Parts subject to fretting. Bolted or pinned connections. b. Rational methods. Methodsmaybeusedwhich donot involve full scale testing but which apply the variables of fatigue strength with acalculation of retirement times in amannerthat provides equivalent reliability to the fatigue testing andsimplified methodsandis acceptable to the Administrator. 10

. TESTINGMETH~DS.

Thefatigue strength of the flight structure may bedetermined in appropriate laboratory tests andevaluated in terms of aloading spectrum. Thestrength indicated bythe test results shouldbereducedby afactor suchthat areplacementtime baseddirectly onthis reducedstrength level andthe loading spectrumof paragraph7will assurethat the probability of failure is extremely remote. This reduction factor shouldbe basedonconsideration of the numberof specimenstested, the variability' of the fatigue results, the effects of service use, and,whereavailable, previous test data for the samematerial or similar componentsaswell as service experience. Thetest methodsoutlined below are considered acceptable. a, S-Nbrves. (1) Fatigue tests shouldbeconductedover arangeof oscillatory stresses or loads to define the S-Ncurves. Fatigue tests shouldbeperfonned at steady stresses or loads representative of those occurring in flight. (2) In order to determine the meanfatigue strength andthe variability in fatigue strength, it is necessaryto test anumberof specimensin establishing SINcurves. In order to accountfor the variability in fatigue strength, areduction factor shouldbeapplied to the meancurve in arriving at aworking SINcurve. This factor should include consideration of the numberof specimenstested,

Par 9 Page 7

### AC20-W

the variability of the fatigue results, and, 

### where w available, previous test data on the samematerial or 

### similar components,aswell as service experiencE. Wherenewmaterials or designs are being evaluated, it is recommended

that a 

### larger reduction factor be 

### useduntil suchtime as additional test data justifying a change are available. The 

### meanand reduced S-N curve should reflect the curve shapes of typical published S-F 

### data on notched or unnotched specimens, as applicable. The reduced S+ curve and the loading spectrum of paragraph 7 should be used in determining replacement

. times. Figure 4 represents the method of constructing

. .a typical S-N curve from the fatigue test data.

1 0 Spectrum'tests. The establishment of replacement times based on fatigue tests in which each specimen is subjected to a spectrum of loading is to include the following considerations: (1) Definition of the test loading spectra based on either: ;(a) Analysis, ,supported by extrapolation of available load history data or prior knowledge where available, or (b) Stress histories based on flight test data obtained

. for flight and ground conditions andmaneuvers cons*idered'appropriate for the particular rotorcraft, and a spectrum allocating percentages of time or frequencies of occurrence to these flight and ground conditions andmaneuvers. (2) Fatigue tests in which the loading spectra are applied such&at effective randomization of loadings is obtained, ((3) Unless performed prior to step '(l), determining by flight test the stress levels associated with each flight condition and maneuver considered appropriate for tlhe particular rotorcraft. (4) Assignment of replacement times. The fatigue test results should be eval.uated in terms of the loading spectrum of .paragraph 7 (if different than test spectr&) 

### andreduced

Pal: IO

### 5/18/76 

## ’ AC 20-95

by a factor considering the number of specimens, the fatigue strength variability, and applicable prior data, in arriving at a replacement time. -_l

C. 'Major system tests. Another method of determining the replacement times is to perform a fatigue test or tests of the m3jor systems, Examples of such testing are whirl tests, tiedown tests, and bench .tests. The test results should be evaluated in terms of the loading spectrum of paragraph 7 (if different from test loading) and reduced by a factor considering number of specimens tested, variability in results, and applicable prior test data in arriving at a replacement time. 11

. COMBINATION

OFREPLACEXENT

TIME ANDFAILSAFE EVAUJATION. It maybe possiLh to extend the replacement time of safe-life components which exhibit limited fail-safe capability by using a combination of the safe-life and fail-safe concepts. This is accomplished by evaluating both the fatigue strength and failsafe characteristics as described elsewhere in this circular and by assigning both a replacemen.t time and inspection period to these components. The replacement time maythen he based on the combined probability of not initiating a fatigue crack at or before the replacement time and the probability that the crack if initiated will be detected prior to catastrophic failure or loss of limit load (or maximumattainable load, whichever is less) carryiq'capability. Theprobability of detection should be based on consideration of the inspection effectiveness, the inspection interval, and the fatigue life remaining after an obvious partial failure. A lower strength reduction factor, commensurate with this probability of detection, maythen be used in the determination of the replacement time. 12l EXTENSIONOFREPLACEXIENT IME. Parts should be replaced or retired 'at the established service period unless additional data indicate that an extension of the service period is justified. Important factors; in the consideration of such extension would be:
a. Recorded load 'data, Recording load data entails instrumenting ' aircraft in service to obtain a representative sampling of actual loads experienced. The data measured should include. airspeed, altitude, and rotor speed versus time, or the airspeed, altitude, and strain ranges versus time, or similar data. The data obtained by instrumenting aircraft in service should provide a basis for correlating the estimated loading spectrum with the actual service experience.

Par 10 Page 9AC20-W 

### 5/18/76

b

. Additional analyses and tests. If test data and analyses based on repeated load tests of additional specimens are obtained, a reevaluation of the initial strength reduction or scatter w.
. factor maybemade.
c. Tests of parts removed from service. Whereconservatism was used in initial calculation of replacement times because of lack of knowledge of service environment, repeated * .?dtests of replaced parts maybe utilized to reevaluate the rni tial scatter factor selected. The tests should closely simulate service loading conditions. dl Rework of the structure. In somecases, rework of the structure may result in an increase in replacement time.

SECTIOK5'. FAIL-SAFE STRENGTH EVALUATION 13l G-L. The fail-safe strength evaluation of the flight structure is intended to insure that, should fatigue cracks initiate, the remaining structure will withstand service loads without failure until the cracks are detected. The fail-safe evaluation generally - encompasses establishing the components which are fail-safe, defining the loading conditions and extent of damagefor which the structure is to be,designed, conducting structural tests and analysis to substantiate that the design objective has been achieved, and establishing inspection programs to assure detection of fatigue damage. Oncomponents predominantly loaded by centrifugal force, care should be taken in selecting limit load to assure that it is the maximumexpected in service. Design features which maybe used in attaining a fail-safe structure are:
a. Selection of materials and stress levels that provide a controlled slow rate of crack propagation combined with high residual strength after injtiation of cracks.

b 0 Design to permit detection of cracks including the use of crack detection systems, in all critical structural elements before &he cracks can becomedangerous or result in appreciable strength loss, and to permit replacement or repair.

c-0. Use of multipath construction and the provision of crack stoppers to limit the growth of cracks. dl Use of composite duplicate structures so that a fatigue crack or failure.occurring in one element of the composite memberwill be

PaEe 10 Par 12

### 5/=/76 AC x1-95

confined to that element and the remaining structure will still possess appreciable load-carrying ability.
e. Use of backup structure wherein onemembercarries all the load, with a second memberavailable and capable of assuming the extra load if the primary memberfails. 140 

### IDENTIFICATION

OFCRITICALPORTIONS

OFFLIGHT

STRUCMRE.Thosepor- w- tions of the flight structure which maybe critical in fatigue should be identified.. Typical portions of the structure are:
a. Rotor blades and attachment fittings. b0 Rotor heads, including hubs, hinges, dampers.
c. Control system components, including control rods, servos, awashplates. d R6cor supporting structure. l

. . . e. Fuselage, including stabilizers and auxiliary lifting surfaces. 150 

### EXTENT

OF

### FAIL-SAFE

DAMAGE.The extent of the partial. failure is to be such that it would be readily detectable during the specified inspection. It may involve complete failure of a principal element, failure 6f more than one element, or only a partiaLfailure of *an.element depending on the rate of crack propagation, the ease of detection, and the inspection interval. Damagein inaccessible areas should extend into inspectable areas. Typical examples of the fatigue damagewhich should be considered are outlined below:
a. Cracks emanating from the edge of structural openings OYcutouts which can be readily detected by visual inspection of the area. b0 A circumferential or longitudinal skin crack In the basic fuselage structure of such a length thatit can be readily detected by a visual 'inspection of the surface area.
c. Complete severance of interior frame elements or stiffeners in addition to avisually detectable crack in the adjacent skin.

Par 13 Page 11AC 20-95

d

. Failure of one element where dual construction is utilized in components.
e. Failure of primary attachments, including control hinges and fittings. 160 DETERMINATIONOFPROBABLECRACK LOCAe~ONS, The probable crack locations are to be determined by tests, analysis, or both. In cases.of unusually critical or complex components or when initial fatigue loadings mayaffect the rate or modeof cracking, the probable crack locations should be determined by fatigue test. V%en determination is madeby analysis, sound engineering judgment should he used and a variety of factors such as the following taken into account:
a. Conducting an analysis to locate areas of low margin of safety. maximumstress and
b l Conducting strain surveys on undamagedstructure to establish points of high stress concentration as well as the magnitude of such concentration.

c. Dcamining static test results to determine locations where excessive deformation occurred. da Determiningfrom fatigue analysis where cracks may initiate.
e. Selecting locations in an element where the stresses in adjacent 'elements would be the maximumwith that element failed. f0 Selecting partial fracture locations in an element wherein high stress concentrations are present in the residual structure.

g a Assessing design detail areas which are prone to fatigue damage based on service experience records of similarly designed components. 17l FAILSAFE DEMONSTRATIO!L It is to be.demonstrated by analysis, tests, or both, that the structure with the partial failures as defined in paragraphs 15and 16 can withstand the maximumload and the repeated loads expected in service during the period prior to detection. The repeated loads should be as defined in the loading spectrum of. paragraph 7 and the structure should be capable of supporting this loading after a partial failure for a sufficient time with respect to the inspection interval to assure that catastrophic

Page 12 Par 15

### 5/33/74 AC 20-95



### failure is extremely remote. Thaloading 

### spectrum should include at least oneapplication of limit load. 

### In test demonstrations, the damage maybe lnit%ated 

### or simulated by cuts madewith afine saw, sharp 

### blade,or guillotiae In those-"caseswhere 

### it is not necessary andnot practical to produce fatigue cracks by tests. 

### In those caseswheredamageis airmrlated at joints or 

### fittings, bolts mayberemovedto simulate failure if this condition wouldbe representative of an actual failure. In 8omeinstances, 

### the failsafe 

### characteristics maybe 

### shownanalytically. Theaaalyelcal approach maybe used 

### whenthe structural configuration involved is essentially similar 

### to onealready verified byfail-safe tests, whether on apreviously 

### approvedtype design, or onother similar areas of the design currently being evaluated. Theanalytical approach mayalso be 

### usedwhen: (1) 

### it canbeshownthat the failure wouldbe detected considerably 

### before the critical crack length is approached; (2) 

### themargins of 

### safety resulting from the analysis are well in excessof the fail-safe residual static strength level; and (3) the 

### stress levels in the partially failed structure and the design are such as 

### to assure-adequate crack propagation time relative to the inspection interval.

18 0 

### TNS~AECTION.

Detection of fatigue cracks before they becomedangerous is the ultimate 

### control in insuring 

### the fail-safe characteristics of the 

### structure. Therefore, the 

### manufacturer shouldprovide sufficient guidance information 

### to assist operators in 

### establishing the frequency and extent of the repeated 

### inspections of 

### the critical structure.

, Par 17 Page 13TABLE1 Percent occurrence I l GROUND

CONDITIONS .. -1.5 

### (a) Racidincrease of r.p.m. on-ground to q@cUzwG clutcbO.5 (b) Taxi'ing with frill cyclic control 0-v 0 

### 5

. (c) Jump

### takeoff --am- (1) (d) Ground-air-ground cycle (1)

. II l HOVERING -- a-w- "-" a---n 

### -2.0 - (a) Steady 

### hovering-'~~~~-n~~n-vnNov-~~~-~---- 0 5 (b) Lateral reversal---------- v-n--o- --o.-- 0 

### 5 (c) Longitudinal reversal------ ~~o~wov~~--- 0

### 5 (d) Rudder reversal- ' -MD-_- ow~-~-vuu 

### 5 III. FORWARD

FLIGHT-POWERON----- 

### '87 S noo~~~~~~--~- 

### 0

. (a) Level 

### flight - 20% VNE-- .

vuv~~nu~~~vn~~o~~~-~~~ 1.0 (b) Level flight - 40% V~------------------------3.0 

### (c) Level 

### flight - 60% VNE----------------------------- 

### 18.0
. (d) 

### Level flight - 80%VW--- 

### ----------..25.0 (e) Maximumlevel 

### flight (but not greater than V&--------15.0

l .

(f) VNE-~~~~-~~-~~-~-~ 

### ~-no~------------~-~~3~0 w 09 (0 (3) (k) (1) ( 1 m 

### ( 1 n ( 1 0 (P) 111% VNE ~~----n~--n a’~~~-~~~~c----~~~- 

### 5 Right turns 

### n30, 60, 90%~--an-n---- (2) 3:o Left t&s 

### - 30, 60, 90%vm------------- ----o----(2) 3.0 -

. Climb (Takeoff power)--------- n--m --- 2.0 Climb (Max. continuous power)--~~~n~~~~.~~~- 

### 4.0 Changeto autorotation from power-on-flight - 

### 30, 60, 90%V&5 Partial power descent (including condition of zero flow through rotor)------- --~--___- ~~~n-~~~-I, 2.0

l Cyclic and 

### collective pull-up from level flight------o(2)

. Pushovers-~~~~~~~-~o~vo~~~~~~~~~---~-- (2) (q) Quick stops~-~-~-~~~~-~~~~---n- aomn--nn,- (2) 

### (r) Flares~-~~~-n-~nw~n-~w-nn ---00-vnv- (2)e (s) Lateral reversals 'at VH---------- -on-n--v- 0 5

. (t) Longitudina1 reversals at VH----------- ---so-- 0 

### 5

. 

### (u) Rudderreversals at VH -v-nnvmn '-n-WV----- 

### 5 (v) Landingapproach--- -~ovo~-no-~-~nnoov~--- 3'0 0 (w) Sidewardflight----. v~v~n~-n--~- -n-n-- 0 5

. 

### (x) Rearwardflight v

~--n-n~n~nv Pm-- v 0 5 IV AIJTOROTATION

- POWER

OFF---------- * -. -vvonnn---
. 9.0 (a) Steady forward flight---------------- -m.--- 

### 2.0 (b) Rapid power recovery 

### from autorotational flight---- 5 (c) Right turns 

### - 30, 60, 90% vm-- -n--n--o-- 

### (2) 1:o (d) Left turns 

### - 30, 60, 90% VNE------------ -n--v- (2) 1.0

Page 14

### S/18/76 AC XL%

(e) Lateral 

### reversals-~----~nn--n----- 0 5 (f) Longitudinal 

### reversals- m-J-n,- --n--o" 0 

### 3 (g) Rudder reversals------ ye--- ~N--v-y' 

### 0 5 (h) 

### Cyclic andcollective pull-ups -rrrnr~-~rr-~rr---- (2) (i) landing approach---------- -ma-am-mwmmn-- 20

(1) We flight every 

### $0minuteswith less frequent rotor stops. 

### (2) A vertical 

### load factor frequency curve should be developed that is representative of the 

### morecritical types of operation. The time spent in each turn should be adjusted to -give 

### the specif led 

### per cent of occurrence.



### Page15AC

### -95 S/18/76

NOTE: DASHED LINES IN FIGURES 1 & 2 INDICATE TEST BOUNDARIES. CROSS-HATCHED AREAS INDICATE OPERATING REGIMES.

VNE 27.1606-,
MAX. DESIGN OR DEMONSTRATED RPM 27.33,27,309 I/

11% VNE 27.33, 27.309

AIRSPEED FIGURE 1. POWER ONROTOR RPM Al RSPEED ENVELOPE 

# I VNE 27.1696 + MAX. DESIGN OR

I DEMONSTRATED RPM 27.33,27.309 ~~---~-~~~ --w-w M


# z a a 0IO a

MIN. DESIGN OR --~~~~~---- DEMONSTRATED RPM

AIRSPEED STEADY STRESS FlGURE3. Oscillatory Stress vs. Steady Stress

STEADY STRESS LEVEL-A 

# I TEST DATA CURVE NO FAILURE AT REQUIRED NUMBER OF CYCLES (6x107)

REDUCED CURVE 1 1 I I

u N I FiGURE 2 POWER OFFROTOR R.P.M. AIRSPEED ENVELOPE FIGUREA Stress vs. Number of Cycles

Page 16_
.. A -,
-*- --.- .-.- -9 F